1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a blade outer air seal and the cooling thereof.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a compressor to deliver a compressed air to a combustor, the combustor combines the compressed air with a fuel to produce a high temperature gas flow, and a turbine that receives the hot gas flow and converts the high temperature flow into mechanical energy to drive a rotor shaft. The efficiency of the gas turbine engine can be improved by increasing the temperature of the flow into the turbine. Prior art turbines include stationary vanes and rotor blades made of high temperature resistant materials in order to maximize the temperature exposure to these parts. Complex cooling circuit are used in the first stage rows of vanes and blades in order to provide cooling such that these parts can be exposed to even higher temperatures that would normally melt the parts.
Another method of increasing the efficiency of the gas turbine engine is to reduce the flow leakage between the rotor blade tips and the shroud casing that forms the blade gap. A plurality of shroud segments that form an annular shroud is fixedly joined to the stator casing and surrounds the rotor blades. The shroud segments are suspended closely atop the blade tips to provide for a small gap or tip clearance between the shroud and the blade tip. In order to reduce the flow leakage across the tip clearance, the tip clearance should be as small as possible to provide for an effective fluid seal during engine operation for minimizing the hot gas flow leakage. However, because the rotor disk and blade have a different thermal expansion and contraction that the casing and shroud segments, the blade tips occasionally rub against the inner surface of the shroud segments and cause abrasion.
The blade tips are directly exposed to the hot gas flow and are difficult to cool properly. The life of the blade is therefore limited because of this difficulty in cooling the tips. Also, when the blade tips rub against the surrounding shroud segments, the blade tips and shroud segments are additionally heated by the friction which also affects the blade useful life. The friction heat generated during a blade tip rub further increases the radial expansion between the tips and the shroud segments, and therefore further increases the severity of the blade tip rub.
Since the shroud segments are also exposed to the hot gas flow through the turbine, the shroud segments are also cooled. Prior art turbine shrouds are cooled by passing cooling air onto the outer surface for impingement cooling to provide backside convective cooling. In addition, film cooling holes are formed in the shroud segments to pass cooling air onto the inner surface of the shroud on which the hot gas flow is exposed. Higher efficiency cooling mechanism such as external film cooling technique has not been widely used in the cooling design. This is primary due to film cooling slots being subject to smear by the passing blade row against the BOAS. Subsequently it loses film cooling capability and shuts off the cooling flow. As a result, over temperature or burn through for the BOAS occurs due to the blade rubbing effect.
Since blade tip rub is unavoidable for maximizing efficiency of the engine, both the turbine shrouds and the blade tips are subject to abrasion wear. U.S. Pat. No. 6,155,778 issued to Lee et al on Dec. 5, 2000 entitled RECESSED TURBINE SHROUD as represented in FIG. 1 discloses a shroud segment used in a gas turbine engine, in which the shroud segments include an inner surface (#50 in this patent) exposed to the hot gas flow, a plurality of recesses (#62 in this patent) opening onto the inner surface 50, and cooling holes to supply cooling air from above the shroud to the recesses 62 to provide film cooling to the shroud inner surface. The recesses 62 are provided for the purpose disclosed in the Lee et al patent for reducing surface area exposed to the blade tips so that during a blade tip rub with the shroud, reduced rubbing of the blade tip with the shroud occurs for correspondingly decreasing frictional heat in the blade tip (see column 3, lines 60-66).
The prior art backside convective cooling used in blade outer air seal (BOAS) cooling design provides cooling to the shroud, but does not provide cooling to the inner shroud surface or the blade tips. Higher efficiency cooling mechanism such as external film cooling has not been widely used in the cooling design. This is primary due to film cooling slots being subject to smear by the passing blade row against the BOAS. Subsequently, it loses film cooling capability and shuts off the cooling flow. As a result, over-temperature or burn out for of the BOAS occurs due to the blade rubbing.
It is therefore an object of the present invention to provide for improved cooling of the shroud segments in a gas turbine engine in order to require less cooling air to provide adequate cooling for the shroud and therefore improve engine efficiency.
It is another object of the present invention to provide for less heat generation due to blade tip to shroud rubbing, and therefore extend the useful life of the rotor blades and shroud segments in the gas turbine engine.
It is another object of the present invention to provide cooling for a BOAS that utilizes both backside multi-impingement compartment cooling and multi-metering plus diffusion cooling for the entire blade outer air seal hot surface.
Another object of the present invention is to provide for a BOAS cooling arrangement in which blade rub will not cause plugging of the cooling holes by the passing blade row against the BOAS.